In order to achieve a high efficiency, a high turbine inlet temperature is required in standard gas turbines. As a result, there arise high NOx emission levels and higher life cycle costs. These problems are mitigated with a sequential combustion cycle, wherein the compressor delivers nearly double the pressure ratio of a conventional one.
The main flow passes the first combustion chamber (e.g. EV combustor), wherein a part of the fuel is combusted. After expanding at the high-pressure turbine stage, the remaining fuel is added and combusted (e.g. SEV combustor). Since the second combustor is fed by expanded exhaust gas of the first combustor, the operating conditions allow self-ignition (spontaneous ignition) of the fuel/air mixture without additional energy being supplied to the mixture (see for example document EP 2 169 314 A2).
Currently convective cooling is used in several combustor parts, e.g. in both the EV and SEV liners. As shown in FIG. 1(a), the cooling air flow 23 of such a combustor part 20 is routed in a cooling channel 22 along the wall 21 to be cooled, and the cooling efficiency can be improved by applying rib turbulators on the wall.
An alternative that can require less cooling air is a combustor part 24 shown in FIG. 1(b) with the application of many small cooling channels 27 (situated between an outer plate 25 and an inner plate 26 of the wall, which channels are situated much closer to the hot side (lower side in FIG. 1). In these channels a higher heat-pick-up can be reached with less cooling mass flow, thus increasing the cooling efficiency. In consequence, less total cooling mass flow is needed, which has a positive impact on the gas turbine performance and emissions.
In the related prior art, several solutions have been proposed with regard to gas turbine combustors:
Document EP 2 295 864 A1 discloses a combustion device for a gas turbine, which shows channels near the wall of the combustion chamber, and which comprises a portion provided with a first and a second wall provided with first passages connecting the zone between the first and second wall to the inner of the combustion device and second passages connecting said zone between the first and second wall to the outer of the combustion device. Between the first and second wall a plurality of chambers are defined, each connected with one first passage and at least one second passage, and defining a Helmholtz damper.
Document U.S. Pat. No. 6,981,358 B2 discloses a reheat combustion system for a gas turbine comprising a mixing tube adapted to be fed by products of a primary combustion zone of the gas turbine and by fuel injected by a lance; a combustion chamber fed by the said mixing tube; and at least one perforated acoustic screen. The acoustic screen is provided inside the mixing tube of the combustion chamber, at a position where it faces, but is spaced from, a perforated wall thereof. In use, the perforated wall experiences impingement cooling as it admits air into the combustion system for onward passage through the perforations of the said acoustic screen, and the acoustic screen damps acoustic pulsations in the mixing tube and combustion chamber.
Document US 2001/0016162 A1 teaches a cooled blade for a gas turbine, in which blade a cooling fluid, preferably cooling air, flows for convective cooling through internal cooling passages located close to the wall and is subsequently deflected for external film cooling through film-cooling holes onto the blade surface, and the fluid flow is directed in at least some of the internal cooling passages in counterflow to the hot-gas flow flowing around the blade, homogeneous cooling in the radial direction is achieved owing to the fact that a plurality of internal cooling passages and film-cooling holes are arranged one above the other in the radial direction in the blade in such a way that the discharge openings of the film-cooling holes in each case lie so as to be offset from the internal cooling passages, in particular lie between the internal cooling passages.
Document WO 2004/035992 A1 discloses a component capable of being cooled, for example a combustion chamber wall segment whereof the walls of the cooling channel include projecting elements of specific shape selectively arranged. The height of the projecting elements ranges between 2% and 5% of the hydraulic diameter of the cooling channel. Thus, the elements are just sufficiently high to generate a turbulent transverse exchange with the central flow in the laminar lower layer, next to the wall, of a cooling flow with fully developed turbulence, thereby considerably enhancing the heat transfer next to the wall of the cooling side without significantly increasing pressure drop in the cooling flow through influence of the central flow.
Document U.S. Pat. No. 5,647,202 teaches a cooled wall part having a plurality of separate convectively cooled longitudinally cooling ducts running near the inner wall and parallel thereto, adjacent longitudinal cooling ducts being connected to one another in each case via intermediate ribs. There is provided at the downstream end of the longitudinal cooling ducts a deflecting device which is connected to at least one backflow cooling duct which is arranged near the outer wall in the wall part and from which a plurality of tubelets extend to the inner wall of the cooled wall part and are arranged in the intermediate ribs branch off. By means of this wall part, the cooling medium can be put to multiple use for cooling (convective, effusion, film cooling).
Document U.S. Pat. No. 6,374,898 B1 discloses a process for producing a casting core which is used for forming within a casting a cavity intended for cooling purposes, through which a cooling medium can be conducted, the casting core having surface regions in which there is incorporated in a specifically selective manner a surface roughness which transfers itself during the casting operation to surface regions enclosing the cavity and leads to an increase in the heat transfer between the cooling medium and the casting.
However, when implementing a near wall cooling channel design on large surfaces, such as for example combustor liners, it is a challenge to assure the feeding and discharging of all near wall channels with cooling air. An example is sketched in FIG. 2: In the gas turbine part 10a of FIG. 2 a feeding channel 12 with an outer channel wall 13a and a separation wall 13 as an inner wall supplies all small cooling channels 15, which run parallel to each other are arranged in a row extending along a predetermined direction, with cooling air. The supplied cooling air 18 enters the feeding channel 12 at one end, enters the cooling channels 15 through their inlets 16, flows through the cooling channels 15, which are embedded in the wall 11 to be cooled, and afterwards, the air enters a discharge channel 14 through cooling channel outlets 17, which discharge channel 14 with its outer wall 13b needs to be separated from the feeding channel 12 by means of the common separation wall 13. From there it is discharged (discharged cooling air 19). On a large surface, e.g. on the liners, several of these elements can be situated next to each other (see FIG. 5).
Since part of the cooling air is fed through each near wall cooling channel 15 (see arrows through the cooling channels in FIG. 2), the remaining cooling mass flow in the feeding channel 12 is decreasing in flow direction. This has a direct impact on the flow velocity and consequently on the static pressure distribution, which is also decreasing along the feeding channel 12. In the discharge channel 14, this effect is reversed: The cooling mass flow and velocity are increasing in flow direction, consequently also increasing the static pressure. Because of these pressure distributions the pressure difference within the near wall channels 15 of one row (from inlet to outlet) is changing along the cooling path and therefore influences the cooling mass flow going through each channel.
However, for a constant cooling performance in all near wall channels it is desirable to have the same mass flow in all channels.